The present invention relates to gas turbine engine fuel nozzles and, more particularly, to main injection structures for gas turbine engine fuel nozzles.
Aircraft gas turbine engines include a combustor in which fuel is burned to input heat to the engine cycle. Typical combustors incorporate one or more fuel injectors whose function is to introduce liquid fuel into an air flow stream so that it can atomize and burn.
Staged combustors have been developed to operate with low pollution, high efficiency, low cost, high engine output, and good engine operability. In a staged combustor, the fuel nozzles of the combustor are operable to selectively inject fuel through two or more discrete stages, each stage being defined by individual fuel flowpaths within the fuel nozzle. For example, the fuel nozzle may include a pilot stage that operates continuously, and a main stage that only operates at higher engine power levels. The fuel flowrate may also be variable within each of the stages.
The main stage includes an annular main injection ring having a plurality of fuel injection ports which discharge fuel through a surrounding centerbody into a swirling mixer airstream. As engine operational requirements become stricter in terms of noise, emissions, and efficiency, there is a need to provide this type of fuel nozzle with greater operational flexibility and control.